The subject matter of the present disclosure relates generally to gas turbine engines. More particularly, the subject matter of the present disclosure relates to an improved system for bleeding air from a gas turbine engine low pressure compressor chamber and a method of making same.
Gas turbine engines such as those used on jet aircraft and industrial gas turbines generally comprise an air inlet, a single or multi-stage compressor chamber, a combustion chamber aft (downstream) of the compressor chamber, a turbine and an exhaust nozzle. Air entering the inlet flows axially through the compressor chamber and into the combustion chamber where it provides oxygen for fuel ignition. As the air passes through the various stages of the compressor its pressure increases. Under certain conditions, such as when the engine is throttled back or during start up, the amount of air required in the combustion chamber is less than that flowing through the compressor chamber. Under these conditions an engine surge may occur, endangering the operation of the aircraft. To mitigate or eliminate engine surges, an air bleed system may be provided within the compression section to temporarily bleed off air and reduce air flow entering the combustion chamber.
Thus there is a need for an air bleed system that includes an annular bleed case with integral structural ligaments that not only provide structural support to the bleed case, but also are aerodynamically designed and oriented with respect to the air flow coming off the compressor rotor to maximize air flow through the bleed ports and reduce pressure loss across the system.
There also is a need for an improved air bleed system that is machined rather than cast for lighter weight and to provide ligaments having a unique geometry and greater strength.
There is also a need for an improved air bleed system that not only directs air out of the engine cavity but also can withstand major engine case loads and support the engine core.